Designers of gas turbine engines for aircraft are constantly trying to increase the thrust-to-weight ratio of their engines. This often results in higher engine operating temperatures and higher stresses in certain engine components, particularly in turbine blades. The combustor temperatures of high-performance gas turbine engines often exceed the melting temperature of the material from which turbine blades are manufactured.
To prevent failure of turbine blades in high-performance gas turbine engines, the turbine blades immediately downstream of the combustor incorporate internal cooling passages through which relatively cool air is passed to cool the blade to prevent the blade temperatures from rising to the combustor temperature. While cooling in this manner is effective at preventing blade failure, inadequate cooling at certain high stress locations of the turbine blade can cause cracks that can ultimately lead to blade failure. One such high stress location is where the trailing edge of the airfoil is joined to the blade platform.
Cooling air for the turbine blade is fed into the turbine blade from below the blade platform, through cooling passages in the blade root. One solution to the problem of cracking at the junction of the trailing edge and the blade platform would be to provide cooling holes in the platform immediately adjacent the trailing edge, and then providing a transverse cooling supply passage through the blade root to connect the cooling holes to the cooling passages in the blade root. Unfortunately, including such a transverse cooling passage in the blade root would weaken the blade while increasing the stress in part of the blade that is already highly stressed.
What is needed is a turbine blade in which the intersection of the airfoil trailing edge and the blade platform is cooled without significantly increasing the stress in the blade during engine operation.